Computer-implemented method for space frame design, space frame construction kit and space frame

ABSTRACT

A computer-implemented method for space frame design involves constructing a load stress map in a geometrical boundary representation of a design space, defining attachment points and load application points in the design space, creating a starting network of interconnecting lines between each two of the attachment points and load application points in the design space, assigning load application factors to each line of the starting network of interconnecting lines based on values of the load stress map, generating potential space frame designs by culling different subsets of lines of the starting network of interconnecting lines for each potential space frame design according to variable culling parameters, evaluating the potential space frame designs with respect to optimization parameters, combining the culling parameters for the potential space frame designs the performance score of which is above a predefined performance threshold, and iterating the steps of generating potential space frame designs and evaluating the potential space frame designs on the basis of the combined culling parameters.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Continuation of co-pending U.S. patent applicationtitled, “COMPUTER-IMPLEMENTED METHOD FOR SPACE FRAME DESIGN, SPACE FRAMECONSTRUCTION KIT AND SPACE FRAME,” filed on Sep. 22, 2017 and havingSer. No. 15/713,572, which is a Divisional of the co-pending U.S. patentapplication titled, “COMPUTER-IMPLEMENTED METHOD FOR SPACE FRAME DESIGN,SPACE FRAME CONSTRUCTION KIT AND SPACE FRAME,” filed on Nov. 30, 2015and having Ser. No. 14/954,052, issued as U.S. Pat. No. 10,430,548. Thesubject matter of these related applications is hereby incorporatedherein by reference.

BACKGROUND OF THE INVENTION Technical Field

The present disclosure relates to a computer-implemented method forspace frame design, a space frame construction kit and a space frameconstructed with a space frame construction kit designed with acomputer-implemented method.

Background

Beams, joists and frames for construction work, for example inaeronautics, civil engineering or architecture, are designed towithstand bending forces acting perpendicular to the direction ofextension of the respective beams. Conventional beams may be implementedas an integral part with flanges at the edges and a web spanning betweenthe flanges. Alternatively, instead of a web, cutter milled struts maybe implemented between parallel running longitudinal support bars, thusleading to decreased weight of the beam due to less material being usedto form the beam.

In order to save weight on board of aircraft, there have been severalattempts to optimize the design of structural aircraft components. Forexample, document WO 2014/111707 A1 discloses a method for designing anobject that includes analysing a digital model corresponding to theobject for portions that have been determined to, during use of theobject, experience relatively high stresses. Those high stress regionsare used to determine which portions of the object are to be producedusing an Additive Manufacturing (AM) process, and which portions of theobject are to be produced using a different suitable process, forexample a machining process. Document DE 10 2010 064 100 A1 discloses apartition wall for separating cabin areas of an aircraft with asandwich-like surface structure.

Structural topology optimization methods for designing structuralobjects according to predefined design criteria have for example beendisclosed in the document WO 2007/076357 A2.

The document U.S. 2009/0224103 A1 discloses a partition wall in anaircraft including a support element composed of individually formedsystem components and a tension-mounted material supported by thesupport element to form an area-shaped partition wall.

There is, however, a need for structural components in aircraft thathave a lower overall weight while at the same time maintainingmechanical stability and the ability to effectively take upstress-induced loads.

SUMMARY OF THE INVENTION

One idea, feature, and/or object of the disclosure herein is thereforeto provide solutions for optimizing the structural topology ofstructural aircraft components in order to decrease the amount ofmaterial needed for building the components.

A first aspect of the disclosure pertains to a computer-implementedmethod for space frame design, the method comprising constructing a loadstress map in a geometrical boundary representation of a design space,defining a plurality of attachment points and load application points inthe design space, creating a starting network of interconnecting linesbetween each two of the plurality of attachment points and loadapplication points in the design space, assigning load applicationfactors to each line of the starting network of interconnecting linesbased on values of the load stress map, generating a plurality ofpotential space frame designs by selectively culling different subsetsof lines of the starting network of interconnecting lines for eachpotential space frame design according to variable culling parameters,evaluating the performance score of each of the plurality of potentialspace frame designs with respect to a number of predefined optimizationparameters, combining the culling parameters for the potential spaceframe designs the performance score of which is above a predefinedperformance threshold, and iterating the steps of generating potentialspace frame designs and evaluating the potential space frame designs onthe basis of the combined culling parameters.

According to a second aspect of the disclosure, a space frameconstruction kit comprises plurality of space frame rods designedaccording to the first aspect of the disclosure, and a plurality ofconnectors for connecting the plurality of space frame rods to a spaceframe.

According to a third aspect of the disclosure, a space frame comprises aplurality of space frame rods designed according to the first aspect ofthe disclosure.

According to a fourth aspect of the disclosure, a structural aircraftcomponent comprises a substantially planar core panel having a spaceframe structure of load bearing space frame rods, with each of the spaceframe rods designed according to the first aspect of the disclosure.

According to a fifth aspect of the disclosure, a computer-readablestorage medium comprises computer-executable instructions which, whenexecuted on a data processing apparatus, cause the data processingapparatus to perform the computer-implemented method according to thefirst aspect of the disclosure.

In the beginning, a two-staged evolutionary route finding algorithm isused for constructing a lightweight space frame structure optimizedtowards high mechanical stability and efficient load transfer anddistribution. In a first stage, an adaptive dynamics scheme forheuristically determining a macroscopic space frame model following themost prominent load paths is employed. The adaptive dynamics scheme ofthe first stage is derived from the adaptive dynamics of a transportnetwork of the amoeboid organism Physarum polycephalum. An optimizationengine varies the input values of a pre-defined parametric model,produces a variety of space frame design options, and discards thelowest performing design options after a performance evaluation with asimplified and therefore rapidly working finite element (FE) model. Theinitial parameters of the surviving designs are used as starting pointfor the evolution of better performing designs, thereby approaching thepareto frontier as stop criterion for the algorithm.

Then, in a second stage, the microstructure of each of the space framemembers determined in the first stage is evolved using a growthalgorithm that iteratively adds member material in each of the spaceframe members in the direction of principal stress. The algorithm mimicsthe way bones and tissues grow in mammal bodies. Maximum strain failuremay here be used as the stop criterion of the iterative optimizationloop.

The generatively designed space frame structure is integrated into asurrounding carrier frame shaping the appearance of the structuralcomponent and adapting it to the surrounding structures in the aircraft.The generative design approach advantageously evaluates a large numberof design options optimizing for both low weight and low structuraldeformation. Thus, this approach is able to reach weight reduction of upto 45% as compared to conventional honeycomb core sandwich structureswhile maintaining equal structural performance.

The designed model is then taken as a basis for manufacturing the spaceframe members in additive manufacturing processes. For greaterflexibility, the space frame members may be divided in sub-componentswith appropriate joint mechanisms. Particularly advantageous mayadditionally be the reduction of costs, weight, lead time, part countand manufacturing complexity coming along with employing any kind oflayer manufacturing technology when designing the components of thespace frame, specifically the space frame rods and/or the concomitantconnectors.

According to an embodiment of the computer-implemented method, theculling parameters may be selected from the group of global linedensity, local line density and line length.

According to a further embodiment of the computer-implemented method,the computer-implemented method may further comprise enriching thestarting network of interconnecting lines with reinforcement linesbetween a node on one of the interconnecting lines and one of theplurality of attachment points and load application points in the designspace or between two nodes on neighboring ones of the interconnectinglines. In some embodiments, the culling parameters may then additionallybe selected from the group of node position on the interconnecting linesand length of the reinforcement lines.

According to a further embodiment of the computer-implemented method,evaluating the performance score of each of the plurality of potentialspace frame designs may be performed using a finite element analysis.

According to a further embodiment of the computer-implemented method,the potential space frame designs may be clustered in a multi-variateoptimization parameter diagram to find space frame designs near thePareto frontier.

According to a further embodiment of the computer-implemented method,the iteration of the steps of generating potential space frame designsand evaluating the potential space frame designs may be terminated whenthe increment in performance score for subsequently generated potentialspace frame designs falls below a termination threshold.

According to a further embodiment of the computer-implemented method,the method may further comprise generating a truss model with amicrostructural framework for each line in the network of lines ofselected ones of the potential space frame designs on the basis ofcorresponding values of the load stress map. In some specificembodiment, the truss models may then be employed as input geometry foran additive manufacturing, AM, process for manufacturing a plurality ofspace frame rods. In some embodiments, the plurality of space frame rodsmay be sub-divided into a number of partial space frame rods having apredefined maximum length.

According to an embodiment of the space frame construction kit, at leasta first one of the space frame rods comprises an angled pin connectorintegrally formed at an end portion of the first space frame rod withangled pins being spaced apart from and protruding parallel to an endface of the first space frame rod, and wherein at least a second one ofthe space frame rods comprises an angled socket connector integrallyformed at an end portion of the second space frame rod with angled tubesas sockets for the angled pins of the first space frame rod, the angledtubes being spaced apart from and protruding parallel to an end face ofthe second space frame rod.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure herein will be explained in greater detail with referenceto exemplary embodiments depicted in the drawings as appended.

The accompanying drawings are included to provide a furtherunderstanding of the present disclosure and are incorporated in andconstitute a part of this specification. The drawings illustrate theembodiments of the present disclosure and together with the descriptionserve to explain the principles of the disclosure herein. Otherembodiments of the present disclosure and many of the intendedadvantages of the present disclosure will be readily appreciated as theybecome better understood by reference to the following detaileddescription. The elements of the drawings are not necessarily to scalerelative to each other. Like reference numerals designate correspondingsimilar parts.

FIG. 1 schematically illustrates an explosion view of components of apartition element of an aircraft according to an embodiment of thedisclosure herein.

FIG. 2 schematically illustrates a front view of a core panel of thepartition element in FIG. 1 according to another embodiment of thedisclosure herein.

FIG. 3 schematically illustrates a computational stage for a digitalsimulation model of the core panel of FIG. 2 according to anotherembodiment of the disclosure herein.

FIG. 4 schematically illustrates a further computational stage for adigital simulation model of the core panel of FIG. 2 according toanother embodiment of the disclosure herein.

FIG. 5 schematically illustrates detail views of connection points ofspace frame rods in a core panel of FIG. 2 according to anotherembodiment of the disclosure herein.

FIG. 6 schematically illustrates a perspective photograph of parts ofthe core panel of FIG. 2 according to another embodiment of thedisclosure herein.

FIG. 7 schematically illustrates a detailed view of the connection typesfor space frame rods of the core panel of FIG. 2 according to evenfurther embodiments of the disclosure herein.

FIG. 8 schematically illustrates stages of a computer-implemented methodfor space frame design according to another embodiment of the disclosureherein.

DETAILED DESCRIPTION

In the figures, like reference numerals denote like or functionally likecomponents, unless indicated otherwise. Any directional terminology like“top”, “bottom”, “left”, “right”, “above”, “below”, “horizontal”,“vertical”, “back”, “front”, and similar terms are merely used forexplanatory purposes and are not intended to delimit the embodiments tothe specific arrangements as shown in the drawings.

Although specific embodiments have been illustrated and describedherein, it will be appreciated by those of ordinary skill in the artthat a variety of alternate and/or equivalent implementations may besubstituted for the specific embodiments shown and described withoutdeparting from the scope of the present disclosure. Generally, thisapplication is intended to cover any adaptations or variations of thespecific embodiments discussed herein.

Some of the components, elements and assemblies as disclosed hereinforthmay be fabricated using free form fabrication (FFF), directmanufacturing (DM), fused deposition modelling (FDM), powder bedprinting (PBP), laminated object manufacturing (LOM), stereolithography(SL), selective laser sintering (SLS), selective laser melting (SLM),selective heat sintering (SHS), electron beam melting (EBM), direct inkwriting (DIW), digital light processing (DLP) and/or additive layermanufacturing (AM). Those techniques belong to a general hierarchy ofadditive manufacturing (AM) methods. Often termed as 3D printing, thosesystems are used for generating three-dimensional objects by creating across-sectional pattern of the object to be formed and forming thethree-dimensional solid object by sequentially building up layers ofmaterial. Any of such procedures will be referred to in the followingdescription as AM or 3D printing without loss of generality. AM or 3Dprinting techniques usually include selectively depositing materiallayer by layer, selectively fusing or solidifying the material andremoving excess material, if needed.

3D or AM techniques may be used in procedures for building upthree-dimensional solid objects based on digital model data. 3D/AMemploys an additive process where layers of material are sequentiallybuilt up in different shapes. 3D/AM is currently used for prototypingand distributed manufacturing with multiple applications in engineering,construction, industrial design, automotive industries and aerospaceindustries.

Space frames within the meaning of the present disclosure may encompassany truss-like structure consisting of or comprising a plurality ofconstruction elements organized in a geometrical assemblage in space.The construction elements are shaped in such a way that forces beingapplied to the assemblage act substantially only on two points of theconstruction elements. The construction elements may themselves have anydesired shape or form, being interconnected to other constructionelements at joints or nodes of the assemblage.

FIG. 1 shows an explosion view of components of a partition element 100of an aircraft in a schematic illustration. The partition element 100may for example as a divider wall between different areas in the cabinof a passenger aircraft. The partition element 100 is exemplary depictedin FIG. 1 as a full-height partition that may for example be installedin an aft area of a passenger aircraft. The partition element 100 mayfor example serve as separating wall between the aft galley and thepassenger compartment. The partition element 100 may be implemented as aline- and retrofit solution using the same interfaces to the airframe asother or conventional partition elements. Of course, the partitionelement 100 serves only as an example for the purposes of descriptionand explanation of some features and aspects of the disclosure herein,and other structural aircraft components may implemented as wellfollowing the very same principles as set out in conjunction with thepartition element 100 of FIG. 1.

The partition element 100 may in some embodiments generally include asubstantially planar core panel 30 having a space frame structure T1 ofload bearing space frame rods RM in the back of the element 100, a coverpanel 20 that is mounted at a front face of the core panel 30, and anattachment panel 10 configured to attach functional elements to thepartition element 100. The attachment panel 10 may for example beadapted to mount a wall-mounted cabin attendant seat bench 5 (CAS bench)with a pivotable seat 4 to the partition element 100. The partitionelement 100 may in particular be designed to comply with FAAairworthiness standards, for example the 16 g dynamic test.

The cover panel 20 may include a substantially rigid protection cover 3with a front face 6 and at least one fabric panel 7 mounted to theprotection cover 3 from its backside, so that the fabric panel 7 issandwiched between the core panel 30 and the protection cover 3 when thecover panel 20 is attached to the core panel 30. As shown in FIG. 1, theprotection cover 3 may in general have an outer shape corresponding tothe outer shape of the core panel 30. One or more cutout apertures 2 ofvarious shapes may be cut out from the material of the protection cover3. The protection cover 3 may for example be made from fibre-reinforcedpolymer material, such as for example glass fibre-reinforced polymer,GFRP, material, natural fibre-reinforced polymer, NFRP, material andcarbon fibre-reinforced polymer, CFRP, material.

The cutout apertures 2 may be matched to the shape of the space framestructure T1 of the core panel 30 in a way that the space frame rods RMare covered from the front side by the cover panel 3 and at least someof the interspaces between neighboring space frame rods RM are visiblethrough the cutout apertures 2 of the cover panel 20.

The fabric panel(s) 7 are mounted to the protection cover 3 andpreferably cover some or all of the cutout apertures 2 from the backsideof the cover panel 20. The fabric panel(s) 7 may for example be fastenedto the protection cover 3 with hook-and-pile fasteners and maythemselves comprise hook-and-pile fasteners on their backside to attachthe fabric panel(s) 7 to the core panel 30. It may be advantageous tohave the outer shape of the fabric panel(s) 7 to match the outer shapeof the respective cutout aperture(s) 2, in a sense that the outer shapeof the fabric panel(s) 7 is congruent to the outer shape of the cutoutaperture(s) 2. The size of the fabric panel(s) 7 may in particular belarger than the size of the corresponding cutout aperture(s) 2, so thata fastening flange portion reaches over the outer rim of the cutoutaperture(s) 2 behind the protection cover 3. Those fastening flangeportions may then be fastened to the protection cover 3 around the rim,for example by aforementioned hook-and-pile fasteners.

A plurality of bores or holes 1 may be drilled or otherwise brought intothe protection cover 3 and the fabric material behind the protectioncover 3, if necessary. The bores 1 may serve as holes for fasteningmembers that may connect the whole cover panel 20 to the core panel 30which may comprise correspondingly positioned bores in the space framerods RM and/or the carrier frame portion as well.

In order to avoid sharpened edges along the inner rim of the cutoutaperture(s) 2, rim protectors 8 may be formed around those rims. The rimprotectors 8 may for example comprises a U-shaped elongated profilemember. Such profile members may be manufactured in AdditiveManufacturing, AM, processes.

The cover panel 20 offers new customization possibilities for airlinesas the cover of the partition is independent from the core panel 30 towhich it is attached. Customers can choose between different coverconcepts such as “closed” flat covers, the flexible integration offabrics, or the integration of light and other features such as screensor displays. Moreover, decorative frames or panel elements may be easilyand replaceably arranged on the smooth front side of the protectioncover, thereby offering vast opportunities for individual airlinebranding and the generation of any desired corporate design.

The core panel 30 as depicted in FIG. 1 is shown in FIG. 2 from thefront in higher detail. Both the core panel 30 and the cover panel 20may have a recessed inner frame at the position of a stretcher flap SF.The stretcher flap SF may be configured to provide access to a stowagearea of a stretcher, for example a stretcher as carriage for stretcherpatients on board of the aircraft. The outer shape of the core panel 30is generally matched to the inner shape of the fuselage section to whichthe core panel 30 may be attached, for example by tie rods attached tohinged interfaces S of the fuselage structure. At the floor portion ofthe cabin, the core panel 30 may for example be fastened to the cabin bybolts at cabin anchor points K. The anchor points K as well as theattachments points of the tie rods for the hinged interfaces S may bereinforced and locally thickened in order to guarantee a smooth andreliable load transfer path into the surrounding fuselage structure.

The core panel 30 of FIG. 2 is generally built up with a macroscopicspace frame structure T1 of a plurality of partially intersecting spaceframe rods R. The particular arrangement of the space frame rod modelsRM may for example be determined by a computer-implemented designprocedure that employs metaheuristic optimization algorithms foroptimizing the load paths through the truss of space frame rods R. Thecore panel 30 may in general comprise a carrier frame F running aroundthe edges of the outer shape of the core panel 30 and the space framestructure T1 extending within the plane spanned by the carrier frame F.The overall thickness of the core panel 30 may in particular be lessthan 3 cm.

All structural members of the core panel 30 may in particular bemanufactured using an Additive Manufacturing, AM, process. Thestructural members of the core panel 30 may for example be made from asuitable material accessible by the AM process, such as for exampleScalmalloy™. Scalmalloy™ is an aluminum-magnesium-scandium alloy(AlMgSc) that has been developed for high and very high-strengthextrusions, offering exceptionally high fatigue properties and the samepositive manufacturing propensities as AlMgSc sheet material. In someembodiments, the core panel 30 may be manufactured in parts so thatsmaller AM machines and systems may be used. For example, it may bepossible to break down the structural topology of the space framestructure T1 into a number of sub-components R (of which two areexemplarily denoted with reference signs in FIG. 2 for illustrativepurposes), such as for example 20 to 150 sub-components R that may beseparately 3D-printed on different ALM systems. Each sub-component R maycontain standardized connectors of different connector types C1 and C2(of which two are exemplarily denoted with reference signs each in FIG.2 for illustrative purposes) which allow for proper connection betweenrespective sub-components R to be joined and for adjustment oftolerances between neighboring sub-components R. In case of damage, theaffected sub-components R may be easily replaced at low cost. Theconnectors of the two different connector types C1 and C2 will beexplained and described in conjunction with FIG. 7 below.

FIG. 6 depicts a perspective photograph of parts of the core panel 30 tobetter illustrate the shape and topology of the space frame assemblageof the core panel 30. As can be seen, the carrier frame on the bottomincludes a thickened anchor point K for connecting the core panel 30 tothe surrounding fuselage structure. Several lower ends of space framerods R are shown to be integrally manufactured with the carrier frame aswell as with each other. Each of the space frame rods R is in itselfformed as a microscopic framework comprising a multitude of laterallyand diagonally running struts which may be interconnected among eachother at local nodes.

The microscopic framework T2 is depicted in conjunction with FIG. 5where three different interconnection node types (A), (B) and (C)between neighboring space frame rods are illustrated exemplarily. Themicroscopic framework T2 may be designed according to local loaddistributions that may be derived from load distribution models underpre-defined boundary conditions. In the node regions N betweenintersecting space frame rods, the laterally and diagonally runningstruts of the microscopic framework T2 of each of the space frame rodsmay be appropriately merged with each other. For each of the rods R, thecore part of the rod may be formed as a truss structure, i.e. astructure consisting of two-force members which are assembled in athree-dimensional structure and connected as nodes. Typically, suchtruss structures may comprise polygonal constructed with straightmembers the ends and sometimes intermediate portions of which areconnected at truss nodes. In the exemplary cases of FIGS. 5 and 6 themicroscopic frameworks T2 take on the shape of a frame having foursubstantially parallel beams extending along the direction of extensionof the space frame rod R and cross-hatched framework patches formed bydiagonally staggered crossbeams between the four substantially parallelbeams.

The topology of the space frame rods R themselves forms the macroscopicframework T1 that may have a generally two-dimensional layout, i.e. thespace frame rods R are substantially lying in one plane of extension (inthe illustrated example of the figures a vertically extending plane).Some or all of the space frame rods R may be equipped with connectors C1or C2 on their respective rod end portions, which connector types areshown in greater detail in FIG. 7.

The connector type C1 may be a rod connector duct RH formed at the endportion of the space frame rods with an inner thread acting as a femaleconnector portion for a male screw connector stud C. The connector studC is formed integrally with an additive manufacturing, AM, process, forexample from Scalmalloy™. At a first shank H1 of the connector stud C aleft-handed outer thread portion is formed, while a second shank H2 thatis located on the opposite side of the connector stud C is formed with aright-handed outer thread portion. If the connector stud C is insertedbetween two neighboring rods R which are both equipped with similar rodconnector ducts RH at their ends, a turning motion D1 will create torquethat causes both thread portions on the first and second shanks H1 andH2 to be simultaneously driven into their respective female threaded rodconnector duct RH. In that way the rods to be joined by the connectortype C1 may be pulled towards each other, with a variable distance inbetween the rods depending on the number of turns applied to theconnector stud C. To aid the application of torque to the connector studC, the connector stud comprises a wrenching contour G that is formedintegrally with the first and second shank H1 and H2 in between bothshanks. The thread formed at the first and second shank H1 and H2 may inparticular have similar pitches. However, in some variants the threadsfor the first and second shank H1 and H2 may have different pitches, forexample if a poka-yoke mechanism is to be put in place to prevent rodsfrom being incorrectly installed in the space frame. The thread startsat the first and second shank H1 and H2 may be single start, however, insome variants a double start may be provided in order to provide moretolerance for assembly.

The connector type C2 may be a half lap splice joint type, where a firstconnector portion RJ1 (as shown in FIGS. 6 and 7) is formed with angledpins being spaced apart from and protruding parallel to an end face ofthe space frame rod. A second connector portion RJ2 (not explicitlyshown in FIG. 6, but in FIG. 7 only) is formed with correspondinglyangled tubes as sockets for the pins being spaced apart from andprotruding parallel to an end face of a space frame rod. The angled pinsand the angled sockets may be brought into interlocking alignment witheach other via a sliding or plugging motion D2 of two corresponding endportions of space frame rods to be connected. There may for example betwo or more than two angled pins and correspondingly angled sockets thatextend in parallel to each other. If there are at least two pins andsockets, the connection C1 is mechanically more stable against torsionalmoments.

The pins and sockets form an undercut in the direction of extension ofthe space frame rods that provides mechanical resistance against thejoint between the space frame rods being pulled apart. In order toprovide more grip between the pins and the inner walls of the sockets,the angled pins may be provided with a knurled outer surface, as may beexemplarily seen in the detail denoted with reference sign “E” in FIG.7. The knurled outer surface may for example comprise a series ofcorrugations, ridges, or diamond patterns that create a large amount ofslight indentations over the pin surface. This enlarges the effectivecontact surface between the outer surface of the angled pins and theinner walls of the angled sockets.

Both connector types C1 and C2 are designed to necessitate only verylittle movement of the space frame rods to be joined relative to eachother. This facilitates connecting neighboring space frame rods R withlittle lateral movement. Particularly in complex structural topologiesof core panels 30, there is generally little leeway for the space framerods R to be displaced with respect to each other in terms of assembly.Thus, the connector types C1 and C2 are favorable connector types whenassembling a complex structural aircraft component.

The space frame T1 may have a generally three-dimensional layout, i.e.for each first plane of extension defined by a subset of space framerods R, another subset of rods R is connected to nodes of the formersubset in a manner that defines at least one further second plane ofextension being arranged under a non-zero angle with respect to thefirst plane of extension. The number of space frame rods R is generallynot limited to any specific number, but instead their number willultimately depend on the result of the optimization algorithm employedto find the optimal design of the space frame T1. Moreover, the number,kind, type and specific design of the connectors at the interconnectionof adjoining space frame rods R may vary depending on the particularoptimized design and/or the desired maximum length of a single spaceframe R.

Generally, a set of space frame rods R and a carrier frame F may form aspace frame construction kit which may be used to construct the desiredstructural component, for example the core panel 30. The space frameconstruction kits as disclosed hereinforth may be used in a lot ofapplications, including—but not limited to—constructions of structuralcomponents in aircraft, interior design, bridge building, vehiclecarriages, civil engineering, applications for children's toys andsimilar. A particular application pertains to the construction of corepanels in structural aircraft components. Such core panels may includespace frame rods for defining an overall outer shape of the structuralaircraft components, for example within a component boundary predefinedby a rigid outer carrier frame.

A computer-implemented method for designing the space frame topology ofthe space frame structure T1 of the core panel 30 will be exemplarilyexplained and described in conjunction with the FIGS. 3 to 5. The methodmay in particular employ metaheuristic optimization algorithms foroptimizing the load paths through the truss of space frame rods R.

At first, existing geometry data is imported into a modelling software.The geometry data may be used to create a boundary representation of thepotential design space, for example a carrier frame model F of thepartition wall of the passenger cabin in an aircraft. Using a solidfinite element analysis model, a load stress map may be calculatedmapped to each point within the boundary representation of the potentialdesign space. Representative loads may be applied in the simulation tostudy the internal load paths and stresses.

As shown in FIG. 3, the modelled carrier frame F may be provided with aplurality of attachment points A at the boundary of the design spacedefining the positions where loads and stress on the space frame isdiverted into the surrounding structures and a plurality of loadapplication points L where major loads are expected to be applied, forexample mounting positions of an attachment panel 10 as shown in FIG. 1.Furthermore, the anchor points K at the carrier frame may be marked upas well in the design space. This definition creates a customizedgeometry as starting settings for the following generative networkoptimization algorithms.

As a starting network, a plurality of interconnecting lines between eachtwo of the plurality of attachment points A and load application pointsL is created in the design space. Additionally, it may be possible toenrich the starting network of interconnecting lines with reinforcementlines. Those reinforcement lines may for example run between nodes N onone of the interconnecting lines and one of the plurality of attachmentpoints A and load application points L in the design space.Alternatively or additionally, other reinforcement lines may run betweentwo nodes N on neighboring ones of the interconnecting lines. Thestarting network thus comprises a much larger number of interconnectingand reinforcement lines than desired for the final space frame design.The starting network of lines (commonly referred to as modelled spaceframe rod lines RM in FIGS. 3 and 4) need then to be culled in order toarrive at a lower number of interconnecting and reinforcement lineswhich run along the expected load paths. Thus, the multi-objectiveoptimization loop involves optimization towards at least minimum weight(corresponding to number of lines) and minimum displacement underpredefined stress.

The multi-objective optimization loop may be subject to similarconsiderations as bionic path-finding models known from the adaptivegrowth of true slime molds. For example, Tero, A.; Kobyashi, R.;Nakagaki, T.; “A mathematical model for adaptive transport network inpath finding by true slime mold”, Journal of theoretical biology No. 244vol. 4, pp. 553-564, Feb. 21, 2007, and Tero, A.; Takagi, S.; Saigusa,T.; Ito, K.; Bebber, D. P.; Fricker, M. D.; Yumiki, K.; Kobayashi, R.;Nakagaki, T.: “Rules for Biologically Inspired Adaptive Network Design”,Science No. 327 vol. 5964, pp. 439-442, Jan. 22, 2010, both disclosecore mechanisms and algorithms for adaptive network formation of trueslime molds captured in a biologically inspired mathematical model.

Each of the interconnecting and/or reinforcement lines is parametrizedwith load application factors. The load application factors are derivedfrom the previously calculated values of the load stress map. Then, thestarting network of interconnecting and/or reinforcement lines isculled. Each culling procedure may run differently with differentparameters defining the culling process, so that a plurality ofdifferent potential space frame designs is generated for furtheranalysis. Culling parameters may for example comprise global linedensity, local line density and line length, or—in case of thereinforcement lines—node position on the interconnecting lines andlength of the reinforcement lines.

The analysis may for example involve the use of an optimization enginethat varies the input values of the parametric model. Each of theobtained potential space frame designs is evaluated with a performancescore that may be obtained using a finite element analysis. Theperformance score may for example take into account a number ofpredefined optimization parameters, such as expected weight of the spaceframe rods manufactured according to the suggested space frame designsand deformation/displacement of the of the space frame rods manufacturedaccording to the suggested space frame designs under external stress.

The potential space frame designs may for example be clustered in amulti-variate optimization parameter diagram to find space frame designsnear the Pareto frontier so that only the potential space frame designswith the optimal performance score for each creation stage is selectedfor further analysis. The culling parameters for the best potentialspace frame designs, i.e. the space frame designs having the bestperformance score are selected for generatively determine a new set ofcombined culling parameters. Depending on the desired convergence speed,only those space frame designs having a performance score above apredefined performance threshold may be selected for this evolutionaryprocedure.

Then, a second generation of potential space frame designs is generatedusing the new set of combined culling parameters is generated which mayagain be subject to the performance evaluation. In this manner, more andmore generations of potential space frame designs may be “evolved” onthe basis of the best properties of the previous generations. Theiteration process may in particular be terminated when the increment inperformance score for subsequently generated potential space framedesigns falls below a termination threshold.

One or several of space frame designs having an optimized macrostructureT1 are then selected for creating a custom generative geometry with anoptimized microstructure T2. To that end, a truss model with amicrostructural framework T2 is generated for each of the suggestedmacroscopic lines in the selected potential space frame designs. FIG. 4exemplarily shows one of the finally selected space frame designs withthe macrostructure T1 on the left side. The detail view on the rightside exemplarily shows the microstructure T2 of the generated trussmodel for one of the modelled space frame rods RM. Again, the generationof the truss model is performed on the basis of corresponding values ofthe pre-calculated load stress map. The generated truss models are thenemployed as input geometry for an additive manufacturing, AM, process.Using the AM process, customized space frame rods may be manufacturedthat fulfil—when combined to a space frame—the desired aims of lowweight, low material consumption and high mechanical stability.

Depending on the AM systems available or desired to be used, some or allof the space frame rods may be sub-divided into a number of partialspace frame rods. Those partial space frame rods may in particular belimited to a predefined maximum length which may correspond to themaximum length that can be manufactured with the available AM systems.The subdivided partial space frame rods may then be equipped with one ofthe connector types C1 and C2 as shown in conjunction with FIG. 7. Forexample, the partial space frame rods may have an end portion formed asfemale-threaded rod connector ducts for connection to a male connectorstud C. Alternatively, the partial space frame rods may be integrallyformed with one of an angled pin connector RJ1 and an angled socketconnector RJ2 at the respective end portions.

FIG. 8 schematically illustrates stages of a computer-implemented methodM for space frame design. The computer-implemented method M may forexample be used to design the space frame for a structural component 100of an aircraft the core panel 30 of which may be manufactured in anadditive manufacturing process with the space frame design as obtainedwith the computer-implemented method M being used as base data for theadditive manufacturing process.

In a first step M1, a load stress map is generated in a geometricalboundary representation of a design space. A second step M2 involvesdefining a plurality of attachment points A and load application pointsL in the design space. Between each two of the plurality of attachmentpoints A and load application points L in the design space a startingnetwork of interconnecting lines may be created in a third step M3.

Optionally, it may be possible to enrich the starting network with aplurality of reinforcement lines that run between a node N on one of theinterconnecting lines and one of the plurality of attachment points Aand load application points L in the design space or between two nodes Non neighboring ones of the interconnecting lines.

In a fourth step M, load application factors are mapped to each line ofthe starting network of interconnecting lines based on values of theload stress map so that a plurality of potential space frame designs maybe generated in a fifth step M5 by selectively culling different subsetsof lines of the starting network of interconnecting lines for eachpotential space frame design according to variable culling parameters.

Each of the plurality of potential space frame designs is evaluated witha performance score in a sixth step M6. This evaluation depends on anumber of predefined optimization parameters. The culling parameters forthe potential space frame designs may be combined in a step M7,specifically for those space frame designs the performance score ofwhich is above a predefined performance threshold. In an evolutionaryalgorithm, the steps M5 of generating potential space frame designs andM6 of evaluating the potential space frame designs are iterated on thebasis of the combined culling parameters.

The space frame construction kit as described as explained above is acheap, extremely light and flexible system that allows for rapidconstruction and deconstruction of multiple structures of varying outerprofile. The space frame construction kit may for example be used tobuild up a core panel of a structural aircraft component, such as forexample an aircraft cabin partition wall. The structural aircraftcomponent built with such a space frame construction kit is easy torepair and thus low in maintenance costs due to single damaged frameworkelements being readily replaceable. Moreover, owing to the modularconstruction of the structural aircraft component, geometry tolerancesof the overall component may be easily compensated for by adjusting theconnectors between the modular parts of the space frame.

A system according to some exemplary embodiments of the presentdisclosure may be provided which includes one or more processingarrangements such as may be found, e.g., in a personal computer orcomputer workstation. Such system can further include a set ofinstructions which are capable of configuring the processing arrangementto perform the exemplary computer-implemented methods described hereinfor designing, constructing, analysing and optimizing models of spaceframes and space frame components. The instructions can be provided on acomputer-accessible medium such as a storage medium. One of skill in theart will recognize that the present disclosure may be implemented as oneor more software processes executable by one or more processors and/orone or more software applications. Additionally, the present disclosureis not described with reference to any particular programming language.It will be appreciated that a variety of programming languages may beused to implement the teachings of the disclosure herein as describedherein. It is also to be understood that the methods may be embodied onany form of memory device or storage medium including all forms ofsequential, pseudo-random, and random access storage devices. Storagemedia as known within the current art include all forms of random accessmemory, magnetic and optical tape, magnetic and optical disks, alongwith various other forms of solid-state mass storage devices, forexample a hard drive, a CD-ROM or DVD-ROM, a tape or floppy disk, aflash drive, or any other solid-state memory storage medium.

In the foregoing detailed description, various features are groupedtogether in one or more examples or examples with the purpose ofstreamlining the disclosure. It is to be understood that the abovedescription is intended to be illustrative, and not restrictive. It isintended to cover all alternatives, modifications and equivalents. Manyother examples will be apparent to one skilled in the art upon reviewingthe above specification.

The subject matter disclosed herein can be implemented in software incombination with hardware and/or firmware. For example, the subjectmatter described herein can be implemented in software executed by aprocessor or processing unit. In one exemplary implementation, thesubject matter described herein can be implemented using a computerreadable medium having stored thereon computer executable instructionsthat when executed by a processor of a computer control the computer toperform steps. Exemplary computer readable mediums suitable forimplementing the subject matter described herein include non-transitorydevices, such as disk memory devices, chip memory devices, programmablelogic devices, and application specific integrated circuits. Inaddition, a computer readable medium that implements the subject matterdescribed herein can be located on a single device or computing platformor can be distributed across multiple devices or computing platforms.

The embodiments were chosen and described in order to best explain theprinciples of the disclosure herein and its practical applications, tothereby enable others skilled in the art to best utilize the disclosureherein and various embodiments with various modifications as are suitedto the particular use contemplated. In the appended claims andthroughout the specification, the terms “including” and “in which” areused as the plain-English equivalents of the respective terms“comprising” and “wherein,” respectively. While at least one exemplaryembodiment of the present invention(s) is disclosed herein, it should beunderstood that modifications, substitutions and alternatives may beapparent to one of ordinary skill in the art and can be made withoutdeparting from the scope of this disclosure. This disclosure is intendedto cover any adaptations or variations of the exemplary embodiment(s).In addition, in this disclosure, the terms “comprise” or “comprising” donot exclude other elements or steps, the terms “a” or “one” do notexclude a plural number, and the term “or” means either or both.Furthermore, characteristics or steps which have been described may alsobe used in combination with other characteristics or steps and in anyorder unless the disclosure or context suggests otherwise. Thisdisclosure hereby incorporates by reference the complete disclosure ofany patent or application from which it claims benefit or priority.

We claim:
 1. A computer-implemented method, comprising: constructing aload stress map in a geometrical boundary representation of a designspace; defining a plurality of attachment points and load applicationpoints in the design space; creating a starting network ofinterconnecting lines between each two of the plurality of attachmentpoints and load application points in the design space; assigning loadapplication factors to each line of the starting network ofinterconnecting lines based on values of the load stress map; generatinga plurality of potential truss structure designs by culling differentsubsets of lines of the starting network of interconnecting lines foreach potential truss structure design according to one or more cullingparameters; evaluating a performance score of each of the plurality ofpotential truss structure designs with respect to a number of predefinedoptimization parameters; and generating at least one truss model basedon the starting network of interconnecting lines and one or more of thepotential truss structure designs selected in accordance with one ormore performance scores.
 2. The computer-implemented method of claim 1,further comprising combining the one or more culling parameters for thepotential truss structure designs having performance scores above apredefined performance threshold.
 3. The computer-implemented method ofclaim 2, further comprising iterating the steps of generating theplurality of potential truss structure designs, and evaluating theperformance score of each of the plurality of potential truss structuredesigns on the basis of the combined culling parameters.
 4. Thecomputer-implemented method of claim 3, wherein the steps of generatingthe plurality of potential truss structure designs and evaluating theperformance score of each of the plurality of potential truss structuredesigns are iterated until an increment in performance score forsubsequently generated potential truss structure designs falls below atermination threshold.
 5. The computer-implemented method of claim 1,wherein the one or more culling parameters comprise at least one of aglobal line density, a local line density, or a line length.
 6. Thecomputer-implemented method of claim 1, further comprising enriching thestarting network of interconnecting lines with reinforcement linesbetween a node on one of the interconnecting lines and one of theplurality of attachment points and load application points in the designspace.
 7. The computer-implemented method of claim 6, wherein the one ormore culling parameters comprise at least one of a node position on theinterconnecting lines or a length of the reinforcement lines.
 8. Thecomputer-implemented method of claim 1, further comprising enriching thestarting network of interconnecting lines with reinforcement linesbetween two nodes on neighboring ones of the interconnecting lines. 9.The computer-implemented method of claim 1, the performance score ofeach of the plurality of potential truss structure designs is evaluatedusing a finite element analysis.
 10. The computer-implemented method ofclaim 1, further comprising clustering the potential truss structuredesigns in a multi-variate optimization parameter diagram to find one ormore truss structure designs near a Pareto frontier.
 11. Thecomputer-implemented method of claim 1, wherein generating the at leastone truss model further comprising generating a truss model with amicrostructural framework for each line in the starting network ofinterconnecting lines of the one or more potential truss structuredesigns on the basis of corresponding values of the load stress map. 12.The computer-implemented method of claim 11, wherein the truss modelsare employed as input geometry for an additive manufacturing process formanufacturing a plurality of truss structure rods.
 13. Thecomputer-implemented method of claim 12, further comprising sub-dividinga plurality of truss structure rods into a number of partial trussstructure rods having a predefined maximum length.
 14. A non-transitorycomputer-readable medium including instructions that, when executed by aprocessor, cause the processor to perform the steps of: constructing aload stress map in a geometrical boundary representation of a designspace; defining a plurality of attachment points and load applicationpoints in the design space; creating a starting network ofinterconnecting lines between each two of the plurality of attachmentpoints and load application points in the design space; assigning loadapplication factors to each line of the starting network ofinterconnecting lines based on values of the load stress map; generatinga plurality of potential truss structure designs by culling differentsubsets of lines of the starting network of interconnecting lines foreach potential truss structure design according to one or more cullingparameters; evaluating a performance score of each of the plurality oftruss structure designs with respect to a number of predefinedoptimization parameters; and generating at least one truss model basedon the starting network of interconnecting lines and one or more of thepotential truss structure designs selected in accordance with one ormore performance scores.
 15. The non-transitory computer-readable mediumof claim 14, further comprising combining the one or more cullingparameters for the potential truss structure designs having performancescores above a predefined performance threshold.
 16. The non-transitorycomputer-readable medium of claim 15, further comprising iterating thesteps of generating the plurality of potential truss structure designs,and evaluating the performance score of each of the plurality ofpotential truss structure designs on the basis of the combined cullingparameters.
 17. The non-transitory computer-readable medium of claim 16,wherein the steps of generating the plurality of potential trussstructure designs and evaluating the performance score of each of theplurality of potential truss structure designs are iterated until anincrement in performance score for subsequently generated potentialtruss structure designs falls below a termination threshold.
 18. Thenon-transitory computer-readable medium of claim 14, wherein the one ormore culling parameters comprise at least one of a global line density,a local line density, or a line length.
 19. The non-transitorycomputer-readable medium of claim 14, further comprising enriching thestarting network of interconnecting lines with reinforcement linesbetween a node on one of the interconnecting lines and one of theplurality of attachment points and load application points in the designspace.
 20. The non-transitory computer-readable medium of claim 19,wherein the one or more culling parameters comprise at least one of anode position on the interconnecting lines or a length of thereinforcement lines.
 21. The non-transitory computer-readable medium ofclaim 14, further comprising enriching the starting network ofinterconnecting lines with reinforcement lines between two nodes onneighboring ones of the interconnecting lines.
 22. The non-transitorycomputer-readable medium of claim 14, the performance score of each ofthe plurality of potential truss structure designs is evaluated using afinite element analysis.
 23. The non-transitory computer-readable mediumof claim 14, further comprising clustering the potential truss structuredesigns in a multi-variate optimization parameter diagram to find one ormore truss structure designs near a Pareto frontier.
 24. Thenon-transitory computer-readable medium of claim 14, wherein generatingthe at least one truss model further comprising generating a truss modelwith a microstructural framework for each line in the starting networkof interconnecting lines of the one or more potential truss structuredesigns on the basis of corresponding values of the load stress map. 25.The non-transitory computer-readable medium of claim 24, wherein thetruss models are employed as input geometry for an additivemanufacturing process for manufacturing a plurality of truss structurerods.
 26. The non-transitory computer-readable medium of claim 25,further comprising sub-dividing a plurality of truss structure rods intoa number of partial truss structure rods having a predefined maximumlength.
 27. A system, comprising: one or more memories storinginstructions; and one or more processors that are coupled to the one ormore memories and, when executing the instructions, are configured toperform the steps of: constructing a load stress map in a geometricalboundary representation of a design space; defining a plurality ofattachment points and load application points in the design space;creating a starting network of interconnecting lines between each two ofthe plurality of attachment points and load application points in thedesign space; assigning load application factors to each line of thestarting network of interconnecting lines based on values of the loadstress map; generating a plurality of potential truss structure designsby culling different subsets of lines of the starting network ofinterconnecting lines for each potential truss structure designaccording to one or more culling parameters; evaluating a performancescore of each of the plurality of truss structure designs with respectto a number of predefined optimization parameters; and generating atleast one truss model based on the starting network of interconnectinglines and one or more of the potential truss structure designs selectedin accordance with one or more performance scores.